The present invention relates generally to combustors in gas turbine engines and, in particular, to a fuel air mixer configured for use in a dome of a gas turbine engine combustor oriented substantially perpendicular to a longitudinal axis through the combustor.
It will be appreciated that emissions are a primary concern in the operation of gas turbine engines, particularly with respect to the impact on the ozone layer by nitrous oxides (NOx), carbon monoxide (CO), and hydrocarbons. In the case of supersonic commercial transport aircraft flying at high altitudes, current subsonic aircraft technology is not applicable given the detrimental effects on the stratospheric ozone. Accordingly, new fuel injection and mixing techniques have been and continue to be developed in order to provide ultra-low NOx at all engine operating conditions.
In response to such emissions concerns, a new combustor has been developed and is discussed in a patent application entitled xe2x80x9cMulti-Stage Radial Axial Gas Turbine Engine Combustor,xe2x80x9dwhich is filed concurrently herewith by the assignee of the present invention, has Ser. No. 09/898,557, and is hereby incorporated by reference. A key component found to provide extremely low levels of NOx at moderate to high power conditions for such aircraft engine was the use of a series of simple mixing tubes as the main fuel injection source. It was found, however, that flame stability and emissions characteristics of a combustor incorporating only such mixing tubes was less capable at low power. Thus, it was determined that an independent pilot fuel injector system would be beneficial for such combustor to improve low power flame stability and meet landing-takeoff (LTO) and idle cycle emissions requirements.
The use of combustion staging has been in practice within the gas turbine engine art for many years to expand the operational range of combustion systems, as well as to provide a broad range of gas turbine power output and applicability. This has typically been accomplished by staging the fuel in a plurality of fuel air mixing devices or modulating the mixing devices independently. In addition, air staging has been performed by having separate and/or isolated annular or cannular combustion zones that can be controlled independently to provide low emissions and a broad range of operation. To date, however, such staging by pilot and main combustion zones has been within substantially the same annular plane.
In light of the foregoing, it would be desirable for a fuel air mixer to be developed which is configured for use in a dome oriented substantially perpendicular to a longitudinal axis through the combustor. It would also be desirable for such fuel a air mixer to be constructed so as to employ a cooling scheme which also improves fuel/air mixing and assists in lowering the fuel-air ratio of the premixture provided to the combustion region of such dome.
In an exemplary embodiment of the invention, a fuel air mixer for a gas turbine engine combustor having a longitudinal axis therethrough is disclosed, wherein the fuel air mixer is configured for use in a dome oriented substantially radial to the longitudinal axis. The fuel air mixer includes a fuel injection assembly having a first end, a second end, a fuel passage extending therethrough, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The fuel air mixer also includes a first end, a second end, a cavity formed in a central portion thereof, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The mixer assembly is configured to receive the fuel injection assembly in the cavity so that the fuel injection assembly and the mixer assembly are able to be connected to an outer casing of the combustor by means of the respective flange portions.